The present invention relates generally to gas turbine engines, and more particularly to internally cooled turbine rotor blades used in such engines.
A gas turbine engine includes a compressor that provides pressurized air to a combustor wherein the air is mixed with fuel and ignited for generating hot combustion gases. These gases flow downstream to one or more turbines that extract energy therefrom to power the compressor and provide useful work such as powering an aircraft in flight. In a turbofan engine, which typically includes a fan placed at the front of the core engine, a high pressure turbine powers the compressor of the core engine. A low pressure turbine is disposed downstream from the high pressure turbine for powering the fan. Each turbine stage commonly includes a stationary turbine nozzle followed in turn by a turbine rotor.
The turbine rotor comprises a row of rotor blades mounted to the perimeter of a rotor disk that rotates about the centerline axis of the engine. Each rotor blade typically includes a shank portion having a dovetail for mounting the blade to the rotor disk and an airfoil that extracts useful work from the hot gases exiting the combustor. A blade platform, formed at the junction of the airfoil and the shank portion, defines the radially inner boundary for the hot gas stream. The turbine nozzles are usually segmented around the circumference thereof to accommodate thermal expansion. Each nozzle segment has one or more nozzle vanes disposed between inner and outer bands for channeling the hot gas stream into the turbine rotor in such a manner that the turbine rotor can do work.
The high pressure turbine components are exposed to extremely high temperature combustion gases. Thus, the turbine blades, nozzle vanes and inner and outer bands typically employ internal cooling to keep their temperatures within certain design limits. The airfoil of a turbine rotor blade, for example, is ordinarily cooled by passing cooling air through an internal circuit. The cooling air normally enters through a passage in the blade""s root and exits through film cooling holes formed in the airfoil surface, thereby producing a thin layer or film of cooling air that protects the airfoil from the hot gases. Known turbine blade cooling circuits often include a plurality of radially oriented passages that are series-connected to produce a serpentine path, thereby increasing cooling effectiveness by extending the length of the coolant flow path.
Similarly, various conventional configurations exist for cooling the nozzle vanes and bands. The most common types of cooling include impingement and film cooling. To effect impingement cooling, the vane airfoil includes one or more perforated hollow inserts that are suitably mounted therein. Cooling air is channeled into the inserts and then impinges against the inner surface of the airfoil for impingement cooling thereof. Film cooling is accomplished by passing the cooling air through film cooling holes formed in the vane airfoil so as to produce a thin layer of cooling air on the outer surface of the vane.
The spaces fore and aft of the rotor disks, commonly referred to as the disk wheel spaces, are in fluid communication with the hot gas stream. Thus, the rotor disks are also subjected to high temperatures, particularly at the disk rim. To prevent overheating of the rotor disks, cooling air is used to purge the fore and aft disk wheel spaces, thereby limiting the ingestion of hot gases.
The cooling air for each of these cooling applications is usually extracted from the compressor. Because the extracted air leads to an associated thermodynamic loss to the engine cycle, it is desirable to keep the amount of air diverted for cooling to a minimum. However, advanced engine designs with increased thrust-to-weight ratios operate at higher turbine inlet temperatures. The higher temperatures require greater overall turbine cooling and make it necessary to cool the blade platform as well. Accordingly, there is a need for improved cooling of turbine components, including the blade platform, without increasing chargeable cooling flow.
The above-mentioned need is met by the present invention which provides a turbine blade including a platform having an internal cavity formed therein and an airfoil extending radially from the platform. An internal cooling circuit is formed in the airfoil for circulating a coolant therethrough, and at least one supply passage extends between the internal cooling circuit and the internal platform cavity for diverting coolant to the internal platform cavity.
The present invention and its advantages over the prior art will become apparent upon reading the following detailed description and the appended claims with reference to the accompanying drawings.